Firing temperatures in modern gas turbine engines continue to increase in response to the demand for higher efficiency engines. Super alloy materials have been developed to withstand the corrosive high temperature environment that exists within a gas turbine engine. However, even super alloy materials are not able to withstand extended exposure to the hot combustion gas of a current generation gas turbine engine without some form of cooling and/or thermal insulation.
Thermal barrier coatings are widely used for protecting various hot gas path components of a gas turbine engine. The reliability of such coatings is critical to the overall reliability of the machine. The design limits of such coatings are primarily determined by laboratory data. However, validation of thermal barrier coating behavior when subjected to the stresses and temperatures of the actual gas turbine environment is essential for a better understanding of the coating limitations. Such real world operating environment data is very difficult to obtain, particularly for components that move during the operation of the engine such as the rotating blades of the turbine. Surface mounted sensors must withstand severe mechanical and thermal loads, and if they become dislodged, they may cause damage to downstream portions of the engine. Furthermore, surface mounted sensors provide information only about conditions that exist at the surface of the thermal barrier coating.
Barrier coatings are also used to protect ceramic matrix composite (CMC) components from the high temperature, oxidizing environment within a gas turbine engine. U.S. Pat. No. 6,197,424 describes an abradable coating for insulating a CMC blade tip seal of a gas turbine. The term “barrier coating” as used herein is meant to include coatings applied to a substrate material to provide at least one of thermal insulation, environmental isolation and abrasion resistance.
U.S. Pat. No. 5,440,300 describes a “smart structure” having sensors and actuators embedded within the material of the structure itself. The patent describes the use of embedded sensors for the detection of stress, strain, vibration, cracks, chemical changes and temperature within the structure. However, the patent does not describe how such sensors may be placed within the structure, and the functionality of the sensors is not intended for the high temperature, corrosive environment of a gas turbine engine.
U.S. Pat. No. 6,000,977 describes a composite structure having electrical leads placed between the plies of material during the fabrication of the structure. The leads are terminated in contact pads at a surface of the structure for making contact with mating pads of a mating component. This structure provides improved interconnectivity for the electrical leads when compared to the previous process of simply allowing the leads to extend out of a trough or bore formed in the structure. These structures and processes are particularly well suited for automotive and aviation applications utilizing layered composite materials. However, they are not useful for instrumenting a component of a gas turbine engine that is covered with a layer of a thermal barrier coating material.